1. Field of the Invention
The present invention relates to a rotor blade for rotorcraft such as a helicopter.
2. Description of the Related Art
FIGS. 13A and 13B are views showing the aerodynamic environment of a helicopter rotor in the forward flight case. As shown in FIG. 13A, when a helicopter 1 having rotor blades of a rotor radius R and rotating at an angular speed .OMEGA. advances at a ground velocity V, an advancing blade in which the ground velocity V is added to a rotor speed .OMEGA.R, and a retreating blade in which the ground velocity V is subtracted from the rotor speed .OMEGA.R are largely different in airspeed from each other.
At the azimuth angle .PSI. (an angle measured counterclockwise from the rearward direction of the helicopter 1) of 90 degrees, the airspeed of the advancing blade reaches the maximum, and the airspeed of the tip of the advancing blade is .OMEGA.R+V. At the position of .PSI.=270 degrees, on the other hand, the airspeed of the retreating blade reaches the minimum, and the airspeed of the tip of the retreating blade is .OMEGA.R-V. The airspeed at an intermediate portion of the blade has a value which is obtained by proportionally distributing .OMEGA.R+V and .OMEGA.R-V. Assuming that .OMEGA.R=795 km/h and V=278 km/h, for example, the airspeed at the position of about 35% from the root end of the retreating blade is zero as shown in FIG. 13A.
When a helicopter flies at high speed, particularly, the airspeed at the tip of the advancing blade is transonic and a strong shock wave is generated. In a drag divergence region shown as a hatched portion in FIG. 13B, this strong shock wave causes an abrupt increase of drag which acts on the blade. Noise which is generated by such a strong shock wave is called high-speed impulsive noise. At this time, in a coordinate system as seen from a rotating rotor blade, a phenomenon which is called delocalization of supersonic region occurs. The generated shock wave propagates to a distant place through a delocalization supersonic region. As a result, high level of noise is observed in the distant place.
In the retreating blade, since the airspeed thereof is significantly lowered, angle of attack .alpha. of the blade must be increased in order to obtain a lift which is equivalent to that of the advancing blade. For this purpose, it is common to carry out pitch control in which a pitch angle of the blade is controlled in accordance with azimuth angle .PSI.. The pitch angle of the blade is controlled using a sinusoidal wave which has a minimum amplitude at .PSI.=90 degrees and a maximum at .PSI.=270 degrees. At this time, as shown in FIG. 13B, the angle of attack a of the blade is changed in a span direction by flapping movement of the blade itself. In the case of .PSI.=90 degrees, for example, the angle of attack a of the blade is about 0 degree at the root, and about 4 degrees at the tip. In the case of .PSI.=270 degrees, the angle of attack .alpha. of the blade is about 0 degrees at the root, and about 16 to 18 degrees at the tip, and exceeds the stall angle of attack. When the angle of attack .alpha. exceeds the stall angle of attack, large changes of lift coefficient Cl and pitching moment coefficient Cm suddenly occur, resulting in that a large vibration of the airframe and fatigue loads in pitch links are generated.
In this way, the high-speed impulsive noise is used as an evaluation item for the advancing blade, and the maximum lift coefficient Clmax and the stall angle of attack are used as evaluation items for the retreating blade. The maximum lift coefficient Clmax is defined by the maximum value of the lift coefficient Cl when the angle of attack .alpha. of a blade having a predetermined aerofoil is gradually increased to reach the stall angle of attack. Usually, as the high-speed impulsive noise and absolute value of the pitching moment coefficient Cm are smaller, or as the maximum lift coefficient Clmax and the stall angle of attack are larger, the blade is judged to be more excellent.
In order to improve the performance of the high-speed flight and to reduce the high-speed impulsive noise, a thin airfoil section may be employed in a blade tip portion. In this method, however, the stall angle of attack is small and also the maximum lift coefficient is small. Therefore, this method is not appropriate. For the above-mentioned purpose, another method in which a simple swept-back angle is formed in the blade tip portion as shown in FIG. 14 is usually employed. A shape provided with the simple swept-back angle is obtained by sweeping back the blade tip portion by a constant swept-back angle. When the simple swept-back angle is formed in the blade tip portion, however, aerodynamic center of the blade tip portion is largely shifted in a rearward direction as shown in FIG. 14. At a position where the aerodynamic center is rearward shifted from a pitch axis by a length .DELTA.X, a moment M about the pitch axis has a value which is obtained by multiplying a lift L with the length .DELTA.X. In the case where the simple swept-back angle is formed as described above, the pitching moment in the direction of the head-down is increased, with the result that control performance is lowered.